Conformal seal bow wave cooling

ABSTRACT

A gas turbine engine includes a combustor. A turbine section is in fluid communication with the combustor. The turbine section includes a first vane stage aft of the combustor. A seal assembly is disposed between the combustor and the first vane stage. The seal assembly includes a plurality of openings communicating cooling airflow into a gap between an aft end of the combustor and the first vane stage. A combustor assembly and method are also disclosed.

BACKGROUND

A gas turbine engine typically includes a fan section, a compressorsection, a combustor section and a turbine section. Air entering thecompressor section is compressed and delivered into the combustionsection where it is mixed with fuel and ignited to generate ahigh-energy exhaust gas flow. The high-energy exhaust gas flow expandsthrough the turbine section to drive the compressor and the fan section.The compressor section typically includes low and high pressurecompressors, and the turbine section includes low and high pressureturbines.

An interface between the combustor exit and the first vane stage canexperience elevated temperatures at localized areas near a leading edgeof each vane. The interface between the combustor exit and the firstvane stage includes a gap. Bow wave phenomena at the leading edge ofeach vane in combination with the gap can result in elevatedtemperatures within and near the gap at this location.

Turbine engine manufacturers continue to seek improvements to engineperformance including improvements to thermal, transfer and propulsiveefficiencies.

SUMMARY

In a featured embodiment, a gas turbine engine includes a combustor. Aturbine section is in fluid communication with the combustor. Theturbine section includes a first vane stage aft of the combustor. A sealassembly is disposed between the combustor and the first vane stage. Theseal assembly includes a plurality of openings communicating coolingairflow into a gap between an aft end of the combustor and the firstvane stage.

In another embodiment according to the previous embodiment, the firstvane stage includes a plurality of vanes with each of the plurality ofvanes including a leading edge and the seal assembly includes aplurality of slots disposed at circumferential positions correspondingwith the leading edge of each of the plurality of vanes.

In another embodiment according to any of the previous embodiments, theplurality of openings extend through the seal assembly and open into acorresponding one of the plurality of slots.

In another embodiment according to any of the previous embodiments, theplurality of openings are disposed in groups spaced circumferentiallyspaced to correspond with the circumferential positions of the pluralityof slots.

In another embodiment according to any of the previous embodiments, theseal assembly includes a radially outer surface and the plurality ofopenings are angled relative to the radially outer surface.

In another embodiment according to any of the previous embodiments, theseal assembly includes aft face that seals against a forward rib of thefirst vane stage.

In another embodiment according to any of the previous embodiments, thecombustor includes a radially outer wall and an aft rib and the sealassembly is disposed on the radially outer wall between the aft rib andforward rib of the first vane stage.

In another embodiment according to any of the previous embodiments, theseal assembly includes an alignment slot that aligns the seal assemblycircumferentially with the first vane stage.

In another featured embodiment, a combustor assembly for a gas turbineengine includes a combustor including an aft end. A seal assemblyextends from the aft end across a gap between the combustor and a firstturbine vane stage. The seal assembly includes a plurality of openingscommunicating cooling airflow into the gap.

In another embodiment according to any of the previous embodiments, theseal assembly includes a plurality of slots disposed at circumferentialpositions corresponding with the leading edge of vanes of the firstturbine stage.

In another embodiment according to any of the previous embodiments, theplurality of openings open into a corresponding one of the plurality ofslots.

In another embodiment according to any of the previous embodiments, theplurality of openings are disposed in groups spaced circumferentiallyspaced to correspond with the circumferential positions of the pluralityof slots.

In another embodiment according to any of the previous embodiments, theseal assembly includes a radially outer surface and the plurality ofopenings are angled relative to the radially outer surface.

In another embodiment according to any of the previous embodiments, thecombustor includes a radially outer wall and an aft rib and the sealassembly is disposed on the radially outer wall between the aft rib anda forward rib of the first vane stage.

In another embodiment according to any of the previous embodiments, theseal assembly includes an alignment slot that aligns the seal assemblycircumferentially with the first vane stage.

In another featured embodiment, a method of cooling a combustor assemblyincludes assembling a seal including a plurality of openings across agap between a combustor and a turbine vane stage aft of the combustor.Cooling air flow is communicated into the gap through the plurality ofopenings in the seal.

In another embodiment according to any of the previous embodiments,includes forming the seal to include a plurality of circumferentialslots and aligning the plurality of circumferential slots with a leadingedge of a turbine vanes within the turbine vane stage.

In another embodiment according to any of the previous embodiments,includes grouping the cooling air holes circumferentially to correspondwith the location of the plurality of circumferential slots.

Although the different examples have the specific components shown inthe illustrations, embodiments of this disclosure are not limited tothose particular combinations. It is possible to use some of thecomponents or features from one of the examples in combination withfeatures or components from another one of the examples.

These and other features disclosed herein can be best understood fromthe following specification and drawings, the following of which is abrief description.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 is a schematic view of an example gas turbine engine.

FIG. 2 is a cross-sectional view of a portion of the gas turbine engine.

FIG. 3 is a front view of a first turbine vane stage.

FIG. 4 is a perspective view of an interface between a combustor and afirst turbine stage.

FIG. 5 is an axial front view of a conformal seal.

FIG. 6 is a front view of a portion of the example conformal seal.

FIG. 7 is an enlarged view of a slot of the example conformal seal.

DETAILED DESCRIPTION

FIG. 1 schematically illustrates a gas turbine engine 20. The gasturbine engine 20 is disclosed herein as a two-spool turbofan thatgenerally incorporates a fan section 22, a compressor section 24, acombustor section 26 and a turbine section 28. Alternative engines mightinclude an augmentor section (not shown) among other systems orfeatures. The fan section 22 drives air along a bypass flow path B in abypass duct defined within a nacelle 18, while the compressor section 24drives air along a core flow path C for compression and communicationinto the combustor section 26 then expansion through the turbine section28. Although depicted as a two-spool turbofan gas turbine engine in thedisclosed non-limiting embodiment, it should be understood that theconcepts described herein are not limited to use with two-spoolturbofans as the teachings may be applied to other types of turbineengines including three-spool architectures.

The exemplary engine 20 generally includes a low speed spool 30 and ahigh speed spool 32 mounted for rotation about an engine centrallongitudinal axis A relative to an engine static structure 36 viaseveral bearing systems 38. It should be understood that various bearingsystems 38 at various locations may alternatively or additionally beprovided, and the location of bearing systems 38 may be varied asappropriate to the application.

The low speed spool 30 generally includes an inner shaft 40 thatinterconnects a fan 42, a first (or low) pressure compressor 44 and afirst (or low) pressure turbine 46. The inner shaft 40 is connected tothe fan 42 through a speed change mechanism, which in exemplary gasturbine engine 20 is illustrated as a geared architecture 48 to drivethe fan 42 at a lower speed than the low speed spool 30. The high speedspool 32 includes an outer shaft 50 that interconnects a second (orhigh) pressure compressor 52 and a second (or high) pressure turbine 54.A combustor 56 is arranged in exemplary gas turbine 20 between the highpressure compressor 52 and the high pressure turbine 54. A mid-turbineframe 58 of the engine static structure 36 is arranged generally betweenthe high pressure turbine 54 and the low pressure turbine 46. Themid-turbine frame 58 further supports bearing systems 38 in the turbinesection 28. The inner shaft 40 and the outer shaft 50 are concentric androtate via bearing systems 38 about the engine central longitudinal axisA which is collinear with their longitudinal axes.

The core airflow is compressed by the low pressure compressor 44 thenthe high pressure compressor 52, mixed and burned with fuel in thecombustor 56, then expanded over the high pressure turbine 54 and lowpressure turbine 46. The mid-turbine frame 58 includes airfoils 60 whichare in the core airflow path C. The turbines 46, 54 rotationally drivethe respective low speed spool 30 and high speed spool 32 in response tothe expansion. It will be appreciated that each of the positions of thefan section 22, compressor section 24, combustor section 26, turbinesection 28, and fan drive gear system 48 may be varied. For example,gear system 48 may be located aft of combustor section 26 or even aft ofturbine section 28, and fan section 22 may be positioned forward or aftof the location of gear system 48.

The engine 20 in one example is a high-bypass geared aircraft engine. Ina further example, the engine 20 bypass ratio is greater than about six(6), with an example embodiment being greater than about ten (10), thegeared architecture 48 is an epicyclic gear train, such as a planetarygear system or other gear system, with a gear reduction ratio of greaterthan about 2.3 and the low pressure turbine 46 has a pressure ratio thatis greater than about five. In one disclosed embodiment, the engine 20bypass ratio is greater than about ten (10:1), the fan diameter issignificantly larger than that of the low pressure compressor 44, andthe low pressure turbine 46 has a pressure ratio that is greater thanabout five 5:1. Low pressure turbine 46 pressure ratio is pressuremeasured prior to inlet of low pressure turbine 46 as related to thepressure at the outlet of the low pressure turbine 46 prior to anexhaust nozzle.

The geared architecture 48 may be an epicycle gear train, such as aplanetary gear system or other gear system, with a gear reduction ratioof greater than about 2.3:1. It should be understood, however, that theabove parameters are only exemplary of one embodiment of a gearedarchitecture engine and that the present disclosure is applicable toother gas turbine engines including direct drive turbofans, land basedturbine engines utilized for power generation as well as turbine enginesfor use in land based vehicles and sea going vessels.

A significant amount of thrust is provided by the bypass flow B due tothe high bypass ratio. The fan section 22 of the engine 20 is designedfor a particular flight condition—typically cruise at about 0.8 Mach andabout 35,000 feet (10,668 meters). The flight condition of 0.8 Mach and35,000 ft (10,668 meters), with the engine at its best fuelconsumption—also known as “bucket cruise Thrust Specific FuelConsumption (‘TSFC’)”—is the industry standard parameter of lbm of fuelbeing burned divided by lbf of thrust the engine produces at thatminimum point. “Low fan pressure ratio” is the pressure ratio across thefan blade alone, without a Fan Exit Guide Vane (“FEGV”) system. The lowfan pressure ratio as disclosed herein according to one non-limitingembodiment is less than about 1.45. “Low corrected fan tip speed” is theactual fan tip speed in ft/sec divided by an industry standardtemperature correction of [(Tram ° R)/(518.7° R)]^(0.5). The “Lowcorrected fan tip speed” as disclosed herein according to onenon-limiting embodiment is less than about 1150 ft/second (350.5meters/second).

The example gas turbine engine includes the fan 42 that comprises in onenon-limiting embodiment less than about twenty-six (26) fan blades. Inanother non-limiting embodiment, the fan section 22 includes less thanabout twenty (20) fan blades. Moreover, in one disclosed embodiment thelow pressure turbine 46 includes no more than about six (6) turbinerotors schematically indicated at 34. In another non-limiting exampleembodiment the low pressure turbine 46 includes about three (3) turbinerotors. A ratio between the number of fan blades 42 and the number oflow pressure turbine rotors is between about 3.3 and about 8.6. Theexample low pressure turbine 46 provides the driving power to rotate thefan section 22 and therefore the relationship between the number ofturbine rotors 34 in the low pressure turbine 46 and the number ofblades 42 in the fan section 22 disclose an example gas turbine engine20 with increased power transfer efficiency.

Referring to FIG. 2 with continued reference to FIG. 1, the examplecombustor 56 includes an axially aft end 62 that is adjacent to anaxially forward face 72 of a first turbine vane stage 64. The firstturbine vane stage 64 includes an upper platform 70 that defines theforward face 72 and has a radially outward extending rib 92. Thecombustor 56 includes a rib 84 that extends radially outward and isspaced apart from the end of the combustor 56.

A conformal seal 76 is disposed between the rib 84 and the forward face72 on a radially outer surface 74 of the combustor 56. The conformalseal 76 extends axially aft from the rib 84 over a radially extendinggap 78 between the combustor 56 and the first turbine vane stage 64.

Referring to FIG. 3 with continued reference to FIG. 2, the firstturbine vane stage 64 includes a plurality of turbine vanes 65 thatextend between the upper platform 70 and a lower platform 75. Each vane65 includes a leading edge 68 facing toward the combustor 56. Theleading edge 68 encounters the high-energy gas flow generated in thecombustor 56 and directs that gas flow into the turbine section 28. Theleading edge 68 of each vane 65 can cause undesired distortions in gasflow that generate non-uniform temperature variations within the gap 78.Bow wave flow phenomena is one such flow distortion that may causeundesired discreet temperature increases. Other flow disruptions thatresult in gas flow entering the gap 78 may also result in undesiredlocalized temperature variations and also will benefit from thisdisclosure.

The example turbine stage 64 includes a plurality of doublets 66 thatare arranged circumferentially about the engine axis A. Each of thedoublets 66 includes two vanes 65 with common upper and lower platforms70, 75. It is within the contemplation of this disclosure to utilizeother turbine vane stage configurations with the disclosed seal 76.

Referring to FIG. 4 with continued reference to FIGS. 2 and 3, thedisclosed example conformal seal 76 includes a plurality of coolingholes 80 that extend from a radially outer surface 96 to a radiallyinner surface 98 that is in communication with the gap 78. Each of thecooling holes 80 are disposed at an angle 100 relative to the radiallyouter surface 96 such that cooling air schematically shown at 102 exitsinto the gap 78. The conformal seals 76 includes a wearing end portion82 that wears down during initial operation to provide a desired sealagainst the axial face 72.

The plurality of cooling air holes 80 are sized to provide a desiredpressure and cooling airflow into the gap 78. In one disclosedembodiment, the plurality of cooling air holes 80 are 0.025 inch (0.635mm) in diameter. The cooling air holes may vary from between 0.015 inch(0.381 mm) and 0.080 inch (2.032 mm) in diameter. It should beunderstood that although an example size of hole is disclosed by way ofexample, other sizes and combinations of cooling hole structures arewithin the contemplation of this disclosure.

Referring to FIGS. 5 and 6 with continued reference to FIGS. 2, 3 and 4,the example conformal seal 76 includes a plurality of slots 86 arrangedcircumferentially about the engine axis A. Each of the slots 86 isaligned with a corresponding leading edge 68 of the vanes 65 within thefirst turbine vane section 64. The cooling holes 80 open on the radialinner surface 98 of the seal 76 within each of the plurality of slots86. The cooling holes 80 communicate cooling airflow to the gap 78 at acircumferential location that corresponds with the leading edge 68 ofeach of the vanes 65.

The plurality of cooling air holes 80 are grouped at the circumferentiallocation that corresponds with the leading edge. In one disclosedembodiment, each grouping includes between 1 and 10 holes. In otherdisclosed embodiment, each grouping of cooling air holes includes 8holes. While specific grouping counts are disclosed, other groupingcounts are within the contemplation of this disclosure.

The conformal seal 76 includes a tab 104 with a slot 106. The slot 106corresponds with slots 95 defined in rib 92 of the vane stage 64 andslot 85 defined as part of the combustor rib 84. An alignment member 94extends through the slots 85, 106 and 95 to align the slots 86 and thecooling holes 80 with the leading edge 68 of each vane 65.

Referring to FIG. 7 with continued reference to FIGS. 4, 5 and 6, eachof the slots 86 provides for communication of cooling airflow 102 intothe gap 78 in a location corresponding with the leading edge 68 of eachof the vanes 65. The cooling holes 80 are grouped circumferentiallyabout the circumference of the conformal seal 76 to correspond with eachof the slots 86.

Cooling airflow 102 is communicated through the conformal seal 76 andinto the gap 78 at the specific circumferential location thatcorresponds with the leading edge 68 of each of the vanes 65.Accordingly, the example conformal seal 76 provides a seal between theend of the combustor and the first turbine stage while also providingdirected cooling airflow to prevent or substantially limit hot gas flowinto the gap 78.

Although an example embodiment has been disclosed, a worker of ordinaryskill in this art would recognize that certain modifications would comewithin the scope of this disclosure. For that reason, the followingclaims should be studied to determine the scope and content of thisdisclosure.

What is claimed is:
 1. A gas turbine engine comprising: a combustorincluding a radially outward extending rib; a turbine section in fluidcommunication with the combustor, the turbine section including a firstvane stage with a forward face aft of the combustor; and a seal disposedbetween an axial surface of the radially extending rib of the combustorand the forward face of the first vane stage, the seal including aforward surface abutting the axial surface of the radially extending ribof the combustor, an aft surface, that is parallel to the forwardsurface, abutting the forward face of the first vane stage, a radiallyouter surface disposed parallel to a radially inner surface, a pluralityof circumferentially spaced apart slots facing the forward face of thefirst vane stage and openings that extend through the seal and aredisposed in groups spaced circumferentially to align with acorresponding one of the circumferentially spaced apart slots, theopenings communicating cooling airflow into a gap between an aft end ofthe combustor and the first vane stage.
 2. The gas turbine engine asrecited in claim 1, wherein the first vane stage includes a plurality ofvanes with each of the plurality of vanes including a leading edge andthe circumferentially spaced apart slots are disposed at circumferentialpositions aligned with a corresponding leading edge of each of theplurality of vanes.
 3. The gas turbine engine as recited in claim 2,wherein the plurality of openings are angled relative to the radiallyouter surface.
 4. The gas turbine engine as recited in claim 1, whereinthe seal assembly includes an alignment slot that aligns the sealassembly circumferentially with the first vane stage.
 5. An assembly fora gas turbine engine comprising: a combustor including an aft end and aradially outward extending rib; and a seal abutting an axial surface ofthe radially outward extending rib from the aft end across a gap betweenthe combustor and a first turbine vane stage, the seal including aforward surface abutting the axial surface of the radially outwardextending rib of the combustor, an aft surface parallel to the forwardsurface and abutting the first turbine vane stage, a radially outersurface disposed parallel to a radially inner surface, a plurality ofcircumferentially spaced apart slots facing a forward face of the firstvane stage and a plurality of openings that extend through the seal andare disposed in groups spaced circumferentially to align with acorresponding one of the circumferentially spaced apart slots forcommunicating cooling airflow into the gap.
 6. The combustor assembly asrecited in claim 5, wherein the plurality of slots disposed atcircumferential positions that align with a leading edge of a respectivevane from a plurality of vanes of the first turbine stage.
 7. Thecombustor assembly as recited in claim 6, wherein the plurality ofopenings are angled relative to the radially outer surface.
 8. Thecombustor assembly as recited in claim 5, wherein the seal includes analignment slot that aligns the seal assembly circumferentially with thefirst vane stage.
 9. A method of cooling an assembly comprising:providing a combustor including an aft end and a radially outwardextending rib; providing a seal abutting an axial surface of theradially outward extending rib from the aft end across a gap between thecombustor and a first turbine vane stage, the seal including a forwardsurface abutting the axial surface of the radially outward extending ribof the combustor, an aft surface parallel to the forward surface andabutting the first turbine vane stage, a radially outer surface disposedparallel to a radially inner surface, a plurality of circumferentiallyspaced apart slots facing a forward face of the first vane stage and aplurality of openings that extend through the seal and are disposed ingroups spaced circumferentially to align with a corresponding one of thecircumferentially spaced apart slots for communicating cooling airflowinto the gap; and communicating cooling air flow into the gap throughthe plurality of openings in the seal.
 10. The method as recited inclaim 9, wherein the plurality of circumferential slots align with aleading edge of a turbine vanes within the turbine vane stage.